Compound aircraft control system and method

ABSTRACT

The Invention is a control system for a compound aircraft. A compound aircraft has features of both a helicopter and a fixed wing aircraft and provides redundant control options. The control system allows an authorized person to select any of plurality of operational objectives each of which is designed to achieve any particular command.

RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.12/168,051, filed Jul. 3, 2008, now issued as U.S. Pat. No. 7,908,044 onMar. 15, 2011. Application Ser. No. 12/168,051 was itself a divisionalapplication from U.S. patent application Ser. No. 11/505,235 by Frank N.Piasecki, et al, filed Aug. 16, 2006, issued as U.S. Pat. No. 7,438,259on Oct. 21, 2008.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The Invention is a control system for a compound aircraft. For trimmedflight, the control system selects a combination of trim controlsettings for the redundant controls of the compound aircraft to achievea pilot's command consistent with a user-selectable objective, such asspeed maximization, fuel consumption minimization, vibration reductionor lifecycle cost reduction. For maneuvering flight, the control systemdistributes control among the redundant control options of the compoundaircraft and may perform that distribution consistent with theuser-selectable objective. The Invention is also a method forcontrolling a compound aircraft.

2. Description of the Related Art

A ‘compound’ aircraft is an aircraft that includes features of bothfixed wing aircraft and rotary wing aircraft. The compound aircraftincludes the elements of a helicopter, including at least one main rotorand a mechanism to overcome the torque response of the rotating mainrotor. The compound aircraft also includes elements of a fixed-wingaircraft, such as a wing. The wing may be equipped with ailerons, flapsor a combination of flaps and ailerons known as ‘flaperons.’ Thecompound aircraft may be equipped with a separate thrust mechanism todrive the aircraft forward, such as a propeller in a ducted fan. Throughthe use of appropriate vanes or sectors that change the configuration ofthe duct, the ducted fan may serve as the mechanism to overcome thetorque response of the rotating rotor blades and to provide yaw control.

A compound aircraft offers several advantages over a conventionalhelicopter. Those advantages include achieving higher flight speeds anddelayed onset of retreating blade stall and leading blade compressioneffects. Although the advantages of a compound helicopter are wellknown, no compound helicopters have been placed in regular operation incommercial or military fleets. One reason is the control complexity ofthe compound aircraft.

The pilot of a conventional helicopter has only limited controls. Thecontrols available for a conventional helicopter having a single mainrotor and a tail rotor are:

Throttle—The pilot can control the amount of power supplied to the rotorblades and to the tail rotor.

Collective pitch—The pilot contemporaneously can change the pitch of allmain rotor blades by an equal amount using the collective pitch control,also known as the ‘collective.’ Contemporaneously changing the pitchangle of all main rotor blades increases or decreases the liftsupporting the helicopter. Increasing the collective and the power willcause the helicopter to rise. Decreasing the collective and the powerwill call the helicopter to sink.

Cyclic pitch—The pilot may use the cyclic pitch control, also known asthe ‘cyclic,’ to cause the pitch angle of the main rotor blades tochange differentially as the main rotor rotates through 360 degrees. Thecyclic pitch control is used to control the pitch and roll of thehelicopter. For example, increasing the pitch angle of a rotor bladewhen the rotor blade is retreating toward the rear of the helicopter anddecreasing the pitch angle when the rotor blade is advancing toward thefront of the helicopter will cause the main rotor plane of rotation totilt forward and hence will cause the helicopter to move forward.

Tail rotor pitch control—For a conventional helicopter having a tailrotor mounted on a boom, a pedal-operated yaw control changes the pitchof the tail rotor blades so that the tail rotor presents more or lessforce countering the torque response of the rotating main rotor. Thepitch of the tail rotor blades therefore controls the yaw of theconventional helicopter.

For a conventional helicopter and for a particular throttle setting,there is only one combination of trim control settings for thecollective, cyclic and tail rotor pitch controls to achieve anyparticular desired trimmed condition of the helicopter. The pilot of theconventional helicopter therefore has few control options.

A compound aircraft will have the aforementioned controls and inaddition will have other controls. For example, the compound aircraftmay feature the following controls:

Flaperon controls—The flaperons (a combination of flaps and ailerons)are located on the wings. When deflected differentially like ailerons,the flaperons may cause the aircraft to roll. When deflected in unisonlike flaps, the flaperons may increase or decrease lift generated by thewing. In hovering flight, the flaperons may be deployed to reduce theeffective wing area and hence reduce the downward force on the wingsfrom the downwash of the main rotor.

Forward thrust control—The compound aircraft may be equipped with aducted fan or other mechanism to provide forward thrust. Thrust providedby the ducted fan or by another mechanism that is not the main rotor isreferred to in this application as “non-rotor forward thrust.”

Rudder/stabililator—The compound aircraft may be equipped with a rudderand with an elevator or stabilator. The rudder controls the yaw of theaircraft, in cooperation with the tail rotor, ducted fan, or othermechanism countering the torque reaction of the rotating main rotor. Anelevator or stabilator controls the pitch of the compound aircraft, incooperation with or instead of the cyclic pitch control.

The pilot of the compound aircraft is presented with a variety ofcontrol combinations to achieve a desired flight condition. For example,if the pilot desires to increase the forward speed of the compoundaircraft, the pilot can increase the non-rotor forward thrust using theforward thrust control, can use the cyclic pitch, stabilator andthrottle controls to pitch the aircraft forward, or can use anycombination of forward thrust control, stabilator, cyclic pitch controland throttle. Each of the possible combinations of trim control settingsoffers advantages and disadvantages. A combination of trim controlsettings that is optimal for one objective (for example, minimizing fuelconsumption) may not be optimal for another objective (for example,minimizing vibration).

Only one combination of trim control settings for the compound aircraftwill be optimal for achieving a particular trimmed condition or forimplementing maneuvering flight commands consonant with also achieving aparticular operational objective. The prior art does not disclose acontrol system for a compound aircraft that allows selection among aplurality of objectives and that then automatically optimizes controlsettings to achieve pilot control commands consistent with the selectedobjective.

BRIEF DESCRIPTION OF THE INVENTION

The Invention is a control system for a compound aircraft. A userselects an overall objective for the control system, such as reducingvibration, increasing performance and speed, reducing lifecycle costs,reducing loading of one or more components, reducing fuel consumption,or any combination of these objectives or of any other desiredobjectives. The overall objective is pre-selected from among a pluralityof overall objectives by the pilot or by another authorized person; forexample, by the owner of the compound aircraft. The control systemreceives a command from a pilot for trimmed flight. The control systemalso receives information from sensors relating to current aircraftcondition (such as attitude, altitude, vertical speed, airspeed, mainrotor speed, control surface positions, acceleration and angular rates).

The control system compares the pilot command to the current aircraftcondition as detected by the sensors and consults a look-up database ofcombinations of trim control settings. The control system applies theuser-selectable overall objective in consulting the look-up database.The control system selects one of the combinations of trim controlsettings for trimmed flight from the look-up database. The selectedcombination of trim control settings provides a control setting for eachof the various control effectors of the compound aircraft to achieve thepilot's intended trimmed flight condition consistent with thepre-selected overall objective. As used in herein, the term ‘controleffector’ means collectively all of the various flight control surfacesand engines of the compound aircraft. The control system applies theselected combination of trim control settings to the control effectorsof the compound aircraft, including the redundant control effectors.

The sensors monitor the current condition of the aircraft and provideconstant feedback to the control system. The control system continuouslyselects and applies different combinations of trim control settings fromthe look-up database as needed to achieve the selected overall objectivefor trimmed flight. If the pilot control inceptors are in ‘detent,’which is a neutral position that does not indicate a commanded change inaircraft condition, a feedback controller regulates the aircraft controleffectors so that the aircraft stays in trim and the selected overallobjective is achieved.

In the event the pilot maneuvers the aircraft, the control system willreceive a pilot command from a control inceptor operated by the pilotand will filter the pilot command using a ‘command filter’ to determinethe commanded change in aircraft condition. The command filterdetermines the dynamic response and thus the handling qualities of thecompound aircraft. The control system compares the filtered pilotcommand to the condition of the aircraft as detected by the sensors andselects a combination of control effector settings to achieve themaneuver.

The control system applies ‘weighting factors’ to control thedistribution of control among the redundant control effectors inmaneuvering flight. The control designer can select a combination ofweighting factors to achieve an overall objective for maneuveringflight, such as minimization of certain structural loads. An authorizedperson, such as the pilot or owner of the aircraft, may select anoverall objective for maneuvering flight from among a plurality ofobjectives. The selected overall objective for maneuvering flight may bethe same as or different from the selected overall objective for trimmedflight. The control system may consult a look-up database and select acombination of weighting factors associated with the selected overallobjective for maneuvering flight. The control system applies theselected combination of weighting factors in allocating control amongthe redundant control surfaces. The control system will supply theselected control settings to the appropriate actuators to achieve themaneuver. When the control inceptors are returned to detent, theaircraft will once again reach trim, with the appropriate controlsettings to achieve the selected overall objective for trimmed flight.

The pilot may change the overall objective for trimmed or maneuveringflight and hence the applied control trim settings or weighting factorsduring flight. For example, the pilot may change the overall objectivefor trimmed flight from ‘reduce vibration’ to ‘maximize speed.’ Thecontrol system then will select a different combination of trim controlsettings to accomplish the new overall objective for trimmed flight.Alternatively, the pilot may not be authorized to change the overallobjective for trimmed or maneuvering flight and the function ofselecting the overall objective may be reserved to another person, suchas the owner of the aircraft.

In an important application of the Invention, a pilot will fly acompound aircraft using only the familiar helicopter flight controls ofcollective pitch, cyclic pitch and tail rotor pitch (pedal yaw control),just as if the pilot were flying a conventional helicopter. The controlsystem receives the collective, cyclic and tail rotor inputs from thepilot and infers the intent of the pilot. The control system thenselects an appropriate combination of trim control settings for thecollective, cyclic, flaperon, forward thrust, elevator, sector, rudderand any other available control effector to best achieve the pilot'sintent, consistent with the pre-selected overall objective for trimmedor maneuvering flight. A pilot skilled in flying a conventionalhelicopter may therefore pilot a compound aircraft using the controlsystem of the Invention and achieve the selected overall objectivewithout simultaneously applying the skills of a fixed-wing pilot.

The control system of the Invention may be a component of a fullyauthorized “fly-by-wire” system in which the control system operates allaircraft flight controls. Alternatively, the control system of theInvention may be configured to operate only a portion of the controls ofthe aircraft. For example, the pilot may directly control thecollective, cyclic pitch and throttle controls, while the control systemof the Invention automatically controls the flaperons and forwardthruster.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a compound aircraft.

FIG. 2 is a side view of a compound aircraft.

FIG. 3 is a rear view of a compound aircraft.

FIG. 4 is a schematic representation of the control system of theinvention.

FIG. 5 is a schematic representation of information flow through thecontrol system of the Invention.

FIG. 6 illustrates the overall control system architecture.

FIG. 7 illustrates the longitudinal/vertical control subsystem.

FIG. 8 illustrates the speed command/speed hold subsystem.

FIG. 9 illustrates the engine torque limiting module

FIG. 10 illustrates the lateral/directional control subsystem.

FIG. 11 illustrates the turn coordination/yaw rate command module.

DESCRIPTION OF AN EMBODIMENT

A. Compound Aircraft Features

The apparatus of the Invention is a control system for a compoundaircraft 2. As shown by FIGS. 1 and 2, the compound aircraft 2 includesfeatures of both a helicopter and a fixed wing aircraft. Those featuresinclude a fuselage 4, a main rotor (or rotating wing) 6, a hub 8 aboutwhich the main rotor 6 rotates and wings 10. Rotation of main rotor 6about hub 8 induces main rotor lift 12. Movement of air across wings 10in response to the forward motion 14 of the compound aircraft 2generates wing lift 16. Rotor lift 12 and wing lift 16 provide lift tothe compound aircraft 2.

Wings 10 feature a wing control surface known as a ‘flaperon’ 18.Flaperon 18 may be moved differentially, in which event flaperons 18 actas ailerons. When used as ailerons, the flaperons 18 in conjunction withwings 10 impart a rolling 8 to fuselage 4. The flaperons 18 also may bemoved in unison, in which event the flaperons 18 act as flaps. When usedas flaps, flaperons 18 change the aerodynamic characteristics of thewing 10 and change wing lift 16.

FIG. 3 is a rear view of the compound aircraft 2. As shown by FIGS. 1, 2and 3, the tail of the compound aircraft 2 features a forward thruster20. Forward thruster 20 is preferably a ducted fan 22. Ducted fan 22features a shroud 24. Shroud 24 improves safety and reduces thelikelihood of damage to the propeller 26 resulting from contact betweenthe propeller 26 and the ground.

Propeller 26 rotates about a ducted fan axis of rotation 28, which isgenerally parallel to the forward direction 14 of compound aircraft 2.Propeller 26 is directly connected to the drive system for main rotor 6,and so the speed of rotation of propeller 26 is directly proportional tothe speed of rotation of main rotor 6 and is not independentlycontrollable. The pitch of propeller 26 is variable, allowing adjustmentof the amount of thrust provided by ducted fan 22.

Sectors 30, shown by FIG. 3, form an adjustable segmented duct toselectably change the direction of thrust of ducted fan 22. Sectors 30,in conjunction with rudder 32, serve to selectably direct the thrust ofducted fan 22 to apply a torque to fuselage 4 contrary to the torqueapplied by main rotor 6. FIG. 3 shows the sectors 30 in a deployedposition and ready to direct ducted fan 22 thrust to counter the torqueof the main rotor 6.

Rudder 32 is adapted to cooperate with sectors 30 to control thedirection of thrust of ducted fan 22. Rudder 32 is in the air stream offan rotor 26 and therefore is capable of affecting yaw of the compoundaircraft 2 at any speed.

Elevator 34 corresponds to the elevator of a fixed wing aircraft.Elevator 34 is in the air stream of fan rotor 26 and is capable ofaffecting the pitch of the compound aircraft 2 at any speed.

B. Control System Overview

FIG. 4 describes the operational relationship between the physicalcomponents of the control system of the Invention. A pilot operatescontrol inceptors 36. The control inceptors 36 correspond to the flightcontrols of a conventional helicopter, with a collective pitch control,a cyclic pitch control, pedal yaw control and throttle. A person skilledat flying a helicopter therefore may operate the compound aircraft 2without simultaneously applying the skills of a fixed-wing pilot. Thecontrol inceptors 36 are connected to a microprocessor 38.

Sensors 40 monitor the condition of the compound aircraft 2 and areconnected to the microprocessor 38. The sensors 40 may monitor compoundaircraft 2 variables such as airspeed, weight and balance parameters,ambient atmospheric conditions, engine torque, propeller 26 torque,vertical speed, pitch rate and attitude, roll rate and attitude, and yawrate.

Microprocessor 38 is operably connected to actuators for each of thecontrol effectors of the compound aircraft 2. Those control actuatorsinclude the cyclic pitch actuator 42, the collective pitch actuator 44,the throttle actuator 46, the first and second flaperon actuators 48,50, the elevator actuator 52, the rudder actuator 54, the sectoractuator 56 and the propeller pitch actuator 58. Each of the actuatorsis adapted by conventional means to operate its associated controleffector under the command of the microprocessor 38.

Computer memory 60 is connected to microprocessor 38. Computer memory 60includes a plurality of selectable overall operational objectives forthe compound aircraft 2. Memory 60 also contains a trim schedulecomprising a look-up database of combinations of trim control settingsselected to achieve each of the selectable overall operationalobjectives for any given condition of the aircraft and any given pilotcommand. The microprocessor 38 consults the database and selects thecombination of trim control settings applicable to the condition of theaircraft, the pilot command, and the selected overall operationalobjective. The microprocessor constantly updates the selection of thecombination of trim control settings based on feedback from the sensorsdetecting the changing aircraft conditions.

The microprocessor applies the selected combination of trim controlsettings for trimmed flight. “Trimmed flight” includes coordinated,level flight and also may include a steady climb or descent or acoordinated turn.

The control system 62 of the Invention uses a “unique trim” concept;that is, when the pilot places the control inceptors in a neutral, ordetent, position, the control system 62 of the compound aircraft 2automatically goes to a trimmed flight condition. The way in which thecompound aircraft 2 is trimmed is selected by the control system 62 ofthe Invention based upon the selected overall operational objective andbased upon the condition of the aircraft 2 as detected by the sensors.

The selected combination of trim control settings defines the controleffector positions and trimmed attitude of the compound aircraft 2required to achieve the optimal trim to accomplish an overalloperational objective. Sensors 40 determine the deviation by thecompound aircraft 2 from the selected combination of trim controlsettings. Feedback paths in the control system 62 can add or subtract tothe effector positions or attitude commands. Since the compound aircraft2 is closed-loop stable, over time if the pilot keeps the controlinceptors in detent, the aircraft 2 will eventually settle into atrimmed condition very close to the optimal trim.

C. Control System Information Flow

FIG. 5 provides an overview of the flow of information through thecontrol system 62 and the principal functions of the control system 62.The control system 62 receives a pilot command 64. The pilot command 64is generated by the control inceptors 36 of FIG. 4 and represents aninstruction by the pilot of the compound aircraft 2. The control system62 also receives a variety of aircraft condition information 66. Theaircraft condition information 66 is generated by sensors 40 as shown byFIG. 4 and provides the microprocessor 38 with the status of thecompound aircraft 2.

The control system 62 examines the sensor information 66 to determinethe condition of the compound aircraft 2 and to evaluate the pilotcommand 64 to determine how the condition of the aircraft will beaffected by the pilot command 64. The control system 62 is fullyauthorized to operate all of the control effectors of the compoundaircraft 2. The control system 62 includes subsystems for selectingoptimal trim 68, speed command/speed holding 70, longitudinal/verticalcontrol 72, lateral/directional control 74 and turn coordination 76. Thecontrol system 62 determines the appropriate subsystem to apply based onthe pilot command 64 and aircraft condition information 66 received. Thecontrol system 62 then applies the protocols of the appropriatesubsystem to determine the appropriate aircraft control settings 78 toaccomplish the pilot command 64 and transmits those control settings 78to the appropriate actuators illustrated by FIG. 4.

The control system 62 relies on feedback to implement the commands ofthe pilot and preferably will incorporate explicit model-followingcontrol architecture, as is known in the art. In such a system, a pilotcommands that the compound aircraft 2 assume a selected flightcondition. The control system 62 determines the condition of theaircraft 2 utilizing sensors 40. The control system 62 determines thechanges to the condition of the aircraft required to reach the commandedcondition. The control system 62 applies an inverse model to determinethe specific control settings 78 required to achieve the commandedcondition and applies those control settings 78 to the control effectorsof the aircraft 2. To compensate for disturbances and modeling orinversion error, the control system 62 measures the changing state ofthe aircraft 2 using the sensors 40 and feeds back the information toupdate the control settings 78 and achieve the commanded condition ofthe aircraft 2.

For a conventional helicopter without the redundant controls of acompound aircraft, the prior art model following/model inversion processis straightforward. The model inversion determines the singlecombination of trim control settings 78 that will achieve the desiredchange in the state of the aircraft and dynamically updates thatcombination of trim control settings 78 to accommodate changingconditions and modeling errors.

For a compound aircraft 2 with redundant controls, the model inversionprocess is more complex. Because of the redundancy, many differentcombinations of trim control settings 78 can achieve a particular changein aircraft state. For any particular change in aircraft state, theforces to achieve that change in state can be allocated among theapplicable control effectors and the control settings adjustedaccordingly.

The control subsystems 68-76 illustrated by FIG. 5 require differentaircraft condition information 66 and require adjustment of differentcombinations of control actuators 42-58, from FIG. 4.

D. Control System Architecture

Each of the control subsystems 68-76 is discussed below and isillustrated in more detail by FIGS. 6-11. The following terms have thefollowing meanings in FIGS. 6-11 and in the discussion below.

β_(p) is the propeller pitch in degrees.

δ_(coll) is the collective control to the mixer in inches.

δ_(e) is the elevator deflection in degrees.

δ_(FO) is the symmetric flaperon deflection in degrees.

δ_(Flat), is the differential elevator deflection in degrees.

δ_(lat) is the lateral control to the mixer.

δ_(long) is the longitudinal control to the mixer.

δ_(yaw) is the yaw control to the VTDP mixer.

φ is the roll attitude in radians.

θ is the pitch attitude in radians.

Ω is the rotor speed in radians/second.

τ_(y) is the yaw response time constant.

a_(y) is lateral acceleration in ft/sec².

‘c’ subscript means ‘post-command filter.’

‘cmd’ subscript means ‘command.’

FADEC means “Fully Automatic Digital Electronic Control. The FADECcontrols fuel to the engine to regulate rotor speed.

GW means gross weight.

HP_(e) is the engine power in standard horsepower.

HP_(VT) the power utilized by the VTDP in standard horsepower.

K_(θ) is the ratio of forward thrust from the rotor to one plus forwardthrust from the propellers.

K_(D) is the derivative gain.

K_(I) is the integral gain.

K_(P) is the proportional gain.

P_(amb) is the ambient pressure in pounds per square inch.

P is the roll rate in radians/second.

q is the pitch rate in radians/second.

r is the yaw rate in radians/second.

s is the Laplace operator.

T_(amb) is the ambient temperature.

τ_(h) is the vertical response time constant.

‘trim’ subscript means ‘optimal trim value.’

U is a pilot command that is filtered to avoid exceedence of operatingparameters.

V is the forward speed in knots or feet/second.

VTDP means ‘vectored thrust ducted propeller’ and is the ducted fan.

V_(Z) is the vertical speed feet/second. Downward is positive.

ω_(nθ) is the pitch response natural frequency.

W is the vertical body velocity in feet/second.

χ_(Bp) is the relationship between propeller pitch and the amount ofthrust generated by the propellers, and varies with airspeed.

χ_(CG) is the longitudinal center of gravity (CG) position.

FIG. 6 is a more detailed diagram of the overall control system 62architecture. Optimum trim schedule 68 is a lookup database ofcombinations of trim control settings 78 and provides control settingsand aircraft attitude for trimmed flight. Optimum trim schedule 68schedules the combinations of trim control settings 78 based on controlinceptor input, aircraft condition information 66 and the selectedoverall operational objective. The aircraft condition information 66 forthe optimum trim schedule 68 may include airspeed, weight and balanceparameters and ambient conditions. The output of the optimum trimschedule 68 comprises optimum trim control settings 78 for all controlsubsystems 70-76. As noted, the overall operational objective isselectable and may include minimizing fuel consumption, minimizingfatigue damage or any other objective or combination of objectives. Asshown by FIG. 6, the optimal rotor speed (Ω_(TRIM)) selected by theoptimal trim schedule subsystem 68 is supplied directly to the FADEC anddetermines engine power output.

As shown by FIG. 6, the pilot manipulates the control inceptors 36 ofcyclic pitch, collective pitch, pedal yaw control and throttle. Thepilot command is processed and the relevant components, as indicate byFIG. 6, are directed to the subsystems of speed command/speed holdcontrol 70, longitudinal/vertical control 72, lateral/directionalcontrol 74 and turn coordination/yaw rate command 76. Each of thesubsystems 70-76 also receives the trim control settings selected by theoptimal trim schedule subsystem 68 and relevant aircraft conditioninformation 66 from the sensors 40. Each of the subsystems 70-76synthesizes the information received and determines control settings toaccomplish the control task assigned to the subsystem, described in moredetail below.

E. Control Mixing

The resulting signals from the subsystems 68-76 must be mixed as shownby FIG. 6. Control mixing is a conventional technique used onrotary-wing and fixed-wing aircraft to minimize cross-coupling effects.Cross-coupling is a change in aircraft 2 attitude or velocity thatresults from the displacement of a pilot control inceptor which is notconsistent with the primary objective of that inceptor. For example, ona helicopter the primary function of the collective lever is to increaseor decreases thrust on the main rotor 6 and thereby change the verticalvelocity of the vehicle. The control system achieves this effect bychanging collective pitch of the main rotor 6 proportional to thecollective lever. However, changes in collective pitch also change thetorque of the rotor 6 and the resultant torque reaction causes theaircraft to yaw. Control mixing can be used to allow tail rotorcollective pitch to also vary with the pilot's collective lever in orderto reduce the collective-to-yaw cross-coupling. Control mixing can beachieved mechanically or in the software of a fly-by-wire controlsystem.

As other examples of control mixing, both the longitudinal/verticalcontrol 72 and the lateral/directional control 74 subsystems control theflaperons 18 to accomplish the different tasks of those subsystems 72,74. The flaperon 18 control signals generated by both subsystems 72, 74must be mixed in the wing mixing module 82 to generate a consistentcontrol signal to the flaperon actuators 48, 50. Similarly, thecollective and cyclic control signals from the speed command/speed holdsubsystem 70, the longitudinal/vertical control subsystem 72 and thelateral/directional control subsystem 74 must be mixed in the rotormixing module 84 prior to sending control settings 78 to the collectiveand cyclic actuators 42, 44.

For a compound aircraft equipped with a Vectored Thrust Ducted Propeller(VTDP) 22, the VTDP 22 provides selectable forward or reverse thrust andyaw control using a combination of rudders 32, sectors 30, and thepropeller 26. The rudder 32 and sectors 30 are in the slipstream of thepropeller 26, so the forces generated by these control surfaces arecoupled to the propeller pitch. The VTDP control mixer 86 is used todetermine the combination of rudder 32, sector 30, and propeller 26pitch to achieve the desired thrust and yaw moment for the given flightcondition and pilot control inputs. In hover and low speed flight, theVTDP 22 is typically configured so that the sectors 30 and rudder 32 arefully deflected and yaw control is achieved using variations inpropeller 26 pitch (this is the “low speed mode”). In forward flight thesectors 30 are retracted, and yaw control is achieved primarily bydeflecting the rudder 32 (“high speed mode”). The VTDP control mixer 86must change the control surface to achieve the low speed mode, highspeed mode, and transitions between the two modes. The inputs to theVTDP control mixer 86 consist of the yaw control input, a propeller 26pitch setting (used for forward thrust control), and calibrated airspeedof the aircraft 2. The yaw control input and propeller 26 setting maycome from the pilot control inceptors, the electronic flight controlsystem, or some combination of the two. The output of the VTDP controlmixer 86 includes the propeller 26 pitch, and the rudder 32 and sector30 deflections. The VTDP control mixer 86 also may include constraintsto ensure proper controllability and to ensure load limits are notexceeded. For example, decreasing propeller 26 pitch setting in highspeed flight may be used to slow down the aircraft 2, but the reversethrust is limited since it reduces flow over the rudder 32 and therebydecreases yaw controllability. There are also airspeed limits on whenthe low speed mode can be used since large deployment of sectors 30 athigh speeds can result in large loads.

F. Longitudinal/Vertical Control

FIG. 7 illustrates the architecture of the longitudinal/vertical controlsubsystem 72. The coupled longitudinal and vertical control subsystem 72controls the longitudinal pitch attitude (θ) and vertical motion (V_(Z))of the compound aircraft 2. The longitudinal/vertical control system 72decreases the pilot workload by selecting control settings 78 to achieveand hold a longitudinal pitch attitude in response to a pilot command64. The longitudinal/vertical control system 72 also will select controlsettings 78 to achieve a pilot command 64 for the aircraft to movevertically and a pilot command 64 to hold a commanded altitude.

The inputs to this subsystem are the commanded pitch attitude, thecommanded vertical speed, and the trim pitch attitude. The aircraftcondition information 66 used by the longitudinal/vertical control 72may include the aircraft pitch rate and attitude, vertical speed andairspeed. The aircraft control settings 78 generated by the controlsystem 62 include longitudinal cyclic, collective, elevator deflectionand symmetric flaperon deflection.

In operation and as shown by FIG. 7, the longitudinal/vertical control72 receives a command for a change in aircraft pitch (Δθ_(CMD)) and achange in vertical speed (ΔV_(ZCMD)). The longitudinal command filter 88will consider the existing pitch rate and pitch attitude and determinesthe desired dynamic response of the aircraft in the pitch and verticalaxes. The longitudinal command filter 88 allows the aircraft to meet theADS-33 Handling Qualities Specifications, if applicable. The pitchresponse is represented by a second order filter. The command filtercalculates the desired pitch attitude, attitude rate, and attitudeacceleration. The desired attitude and attitude rate are compared tomeasured values and multiplied by gains and summed with the desiredpitch acceleration. The summed value is the pitch pseudo-control, whichrepresents the commanded pitch attitude acceleration. The modifiedcommand is indicated as U_(θ) on FIG. 7.

As shown by FIG. 7, the pilot command 64 for a change in vertical speedis treated similarly. The vertical response is represented by afirst-order filter. The time constant can be selected to meet ADS-33heave response specifications. The outputs of the filter are the desiredvertical speed and vertical acceleration. The desired vertical speed iscompared to the measured value and passed through a proportional plusintegral compensator. This is summed with the desired verticalacceleration to calculate the vertical axis pseudo-control, whichrepresents the commanded vertical acceleration of the aircraft.

From FIG. 7, the pitch and vertical axis pseudo-controls are passedthough a model inversion. The inversion module 92 determines whatchanges to the control settings (δ) must be made to achieve the filteredcommand (U). The inversion module 92 considers the trim settings fromthe optimum trim schedule 68 and the condition of the aircraft from thesensors 40. The inversion module 92 is ‘airspeed scheduled’ because theresults of the model vary with air speed. The inversion module 92allocates the pitch commands between the redundant controls that affectthe pitch of the compound aircraft 2; namely, the longitudinal cycliccontrol and the elevator 34 deflection control. For example, theinversion module 92 may allocate more of the pitch control duties to theelevator 34 and less to the longitudinal cyclic to reduce the loads onthe rotor hub 8.

To implement the model inversion of the longitudinal/vertical control72, the pitch pseudo-control (which represents the second derivative ofthe pitch Euler angle) and the vertical pseudo-control (which representsthe vertical acceleration in the inertial frame) are converted to pitchacceleration and vertical acceleration in the body axes. Also, thedesired pitch attitude rate and vertical speed are converted to bodyaxes.

$\begin{matrix}\begin{matrix}{{\overset{.}{q}}_{c} = \frac{U_{\theta}}{\cos\;\phi}} & {{\overset{.}{w}}_{c} = \frac{U_{V_{z}} + {u\;\cos\;\theta{\overset{.}{\theta}}_{c}}}{\cos\;{\theta cos\phi}}} \\{q_{c} = \frac{{\overset{.}{\theta}}_{c}}{\cos\;\phi}} & {w_{c} = \frac{V_{zc} + {u\;\sin\;\theta}}{\cos\;{\theta cos\phi}}}\end{matrix} & {{Eqn}.\mspace{14mu} 1}\end{matrix}$At any given airspeed the linearized short-period longitudinal dynamicscan be represented by:

$\begin{matrix}{\begin{bmatrix}{\Delta\;\overset{.}{w}} \\\overset{.}{q}\end{bmatrix} = {\begin{bmatrix}Z_{w} & {Z_{q} + u_{0}} \\M_{w} & M_{q}\end{bmatrix}{\quad{\begin{bmatrix}{\Delta\; w} \\q\end{bmatrix} + {\begin{bmatrix}Z_{\delta_{long}} & Z_{\delta_{coll}} & Z_{\delta_{e}} & Z_{\delta_{F\; 0}} \\M_{\delta_{long}} & M_{\delta_{coll}} & M_{\delta_{e}} & M_{\delta_{F\; 0}}\end{bmatrix}\begin{bmatrix}{\Delta\delta}_{long} \\{\Delta\delta}_{coll} \\{\Delta\delta}_{e} \\{\Delta\delta}_{F\; 0}\end{bmatrix}}}}}} & {{Eqn}.\mspace{14mu} 2}\end{matrix}$This represents a linear state space model of the form:{dot over (x)}=Ax+Bu  Eqn. 3The B matrix in this case is wide due to the redundant controlseffectors. Normally, model inversion is achieved by taking the inverseof B:u=B ⁻¹({dot over (x)} _(des) −Ax _(des))  Eqn. 4However, in this case the matrix is square and cannot be inverted. Infact, since there are redundant controls, there are many differentcombinations of controls that will achieve the desired pitch andvertical body accelerations. One possible solutions is to use a leftinverse of B,B ^(L) =B ^(T)(BB ^(T))⁻¹  Eqn. 5which results in the control vector u with minimum norm. However, onemight want to put different weighting on the magnitudes of each of thedifferent control effectors. A weighted left inverse of the B matrix canbe represented as:B ⁺ =W(BW)^(T)[(BW)(BW)^(T)]⁻¹  Eqn. 6the control law can then be represented by:u=B ⁺({dot over (x)} _(des) −Ax _(des))  Eqn. 7which gives the a control vector that achieves the desired accelerationswhile minimizing the norm of the vector Wu.

In general, the longitudinal dynamics of the aircraft will varysignificantly with airspeed. An airspeed scheduled model of thelongitudinal dynamics can be represented by:

$\begin{matrix}{\begin{bmatrix}{\Delta\;\overset{.}{w}} \\\overset{.}{q}\end{bmatrix} = {{{A_{lon}(V)}\begin{bmatrix}{\Delta\; w} \\q\end{bmatrix}} + {{B_{lon}(V)}\begin{bmatrix}{\Delta\delta}_{long} \\{\Delta\delta}_{coll} \\{\Delta\delta}_{e} \\{\Delta\delta}_{F\; 0}\end{bmatrix}}}} & {{Eqn}.\mspace{14mu} 8}\end{matrix}$The control law is given by:

$\begin{matrix}{{\begin{bmatrix}{\Delta\delta}_{long} \\{\Delta\delta}_{coll} \\{\Delta\delta}_{e} \\{\Delta\delta}_{F\; 0}\end{bmatrix} = {{B_{lon}^{+}(V)}\left( {\begin{bmatrix}{\;{\overset{.}{w}}_{c}} \\{\overset{.}{q}}_{c}\end{bmatrix} - {{A_{lon}(V)}\begin{bmatrix}{w_{c} - w_{trim}} \\q_{c}\end{bmatrix}}} \right)}}{{{where}\mspace{14mu} w_{trim}} = {V\;\tan\;\theta_{trim}}}} & {{Eqn}.\mspace{14mu} 9}\end{matrix}$The weighted left inverse is defined by:

$\begin{matrix}{{B^{+} = {{W_{lon}\left( {B_{lon}W_{lon}} \right)}^{T}\left( {B_{lon}W_{lon}W_{lon}^{T}B_{lon}^{T}} \right)^{- 1}}}{W_{lon} = \begin{bmatrix}w_{long} & \; & \; & \; \\\; & w_{coll} & \; & \; \\\; & \; & w_{e} & \; \\\; & \; & \; & w_{\delta_{F}}\end{bmatrix}}} & {{Eqn}.\mspace{14mu} 10}\end{matrix}$

The parameters w_(long), w_(coll), w_(e), and w_(δ) _(F) , are selectedto get the desired distribution of control to the longitudinal cyclic,collective, elevator, and flaperons respectively. For example, the w_(e)term can be increased and the w_(long) term increased to select moreelevator relative to longitudinal cyclic when maneuvering the compoundaircraft 2 in pitch. The distribution of the control among the redundanteffectors may be scheduled, stored in computer memory 60 and selected bythe microprocessor 38 to achieve user-selected overall operationalobjectives, such as limiting stress on a particular part or subsystem,minimizing vibration, reducing lifecycle costs, maximizing fuel economy,maximizing speed, or any other operational objective.

The required changes to the control settings 78 from theairspeed-scheduled inversion module 92 are directed to the three mixers82-86, as described above relating to FIG. 6.

G. Speed Command/Speed Hold

FIG. 8 illustrates the operation of the speed command/speed holdsubsystem 70. The speed command/speed hold system 70 allows the controlsystem 62 to respond to longitudinal acceleration commands ({dot over(V)}_(CMD)) and holds the existing forward speed (V_(Z)) of the aircraftwhen no change in command is received. The overall purpose of the speedcommand/speed hold system 70 is to simplify the pilot task ofcontrolling forward speed. In a helicopter, the forward acceleration anddeceleration, and hence forward speed, is controlled by the pitchattitude (θ) of the main rotor disk. In the compound aircraft 2, theforward speed of the aircraft 2 is controlled not only by pitch attitudeof the main rotor disk but also by the propeller blade pitch (β_(P)).The speed command/speed hold 70 control subsystem integrates thepropeller blade pitch control with the pitch attitude of the aircraft 2so that the pilot can control forward speed with a single inceptor.

The engine torque limiting module 94, illustrated by FIG. 8 and shown indetail in FIG. 9, constrains the pilot command 64 for forwardacceleration or deceleration by imposing engine torque (power) limits.Engine torque limits prevent the control system 62 from calling forcontrol settings 78 that require more than 100% or less than 0% of theavailable output of the engine. The aircraft condition information 66required by the torque limiting module include the pilot command 64 forforward acceleration ({dot over (V)}_(CMD)), airspeed, engine torque,propeller blade torque and ambient conditions. The output of the torquelimiting module is a constrained acceleration ({dot over(V)}_(CMD))_(LIM).

The constrained acceleration command is integrated to create a commandforward speed. This is compared to the measured forward speed tocalculate an error signal which is passed through a PI compensator andthe resulting signal is added to the commanded acceleration to create apseudo-control for forward speed. The PI compensator provides the SpeedHold function of the controller.

The forward speed pseudo-control represents the desired forwardacceleration. A simplified linear model of the speed dynamics can berepresented as:{dot over (V)}=X _(β) _(p) Δβ_(p) −gΔθ  Eqn. 11

Acceleration is proportional both to main rotor disk pitch attitude andthe change in propeller pitch. The equation can be inverted to calculatethe change in pitch attitude and propeller pitch needed to achieve thedesired forward acceleration:

$\begin{matrix}{{{\Delta\theta}_{cmd} = {{- \frac{K_{\theta}}{g}}U_{\overset{.}{V}}}}{{\Delta\beta}_{p} = {\frac{\left( {1 - K_{\theta}} \right)}{X_{\beta_{p}}}U_{\overset{.}{V}}}}} & {{Eqn}.\mspace{14mu} 12}\end{matrix}$

The X_(β) _(p) term represents the sensitivity of auxiliary thrust dueto propeller pitch changes. It is a function of flight condition, and inthis case is scheduled with airspeed. It could be scheduled with otherparameters if necessary, but an exact value of the propeller sensitivityis not needed.

The change in propeller pitch is added to the trim propeller pitchspecified by the optimal trim schedule. Likewise the Δθ_(cmd) term isadded to the optimal trim pitch attitude in the Longitudinal AFCS. Thepropeller pitch setting can be limited to observe torque limits on thepropeller gearbox using a similar scheme used for engine torquelimiting.

The gain K□ represents the amount of pitch attitude used foracceleration relative to auxiliary thrust. So if K□□=1 the aircraftaccelerates like a helicopter using changes in pitch attitude and thepropeller pitch only changes to follow the optimal trim schedule. IfK□□=0, the aircraft accelerates using auxiliary thrust while the pitchattitude follows the optimal trim schedule. The designer can select K□□□to achieve optimal distribution of control in maneuvering flight. Thegain may be scheduled with current aircraft condition and selectedoverall operational objective.

The propeller sensitivity schedule 96 of FIG. 8 represents therelationship between propeller pitch and the amount of thrust generatedby the propeller 26. This relationship is referred to by the term X_(βP)and is a function of airspeed.

The speed command/speed hold subsystem 70 determines the changes to thepropeller pitch (β_(P)) and to the pitch attitude (θ) of the aircraftnecessary to achieve the allocated commanded forward acceleration ordeceleration. If a compound aircraft 2 is accelerating from the minimumpower speed or is descending, it can accelerate very quickly becauseexcess power is available. If the compound aircraft 2 is climbing,operating near the maximum speed or performing an aggressive turn, thenavailable excess power and hence acceleration is limited. The maximumacceleration may even be negative because the pilot needs to bleed offairspeed to perform a maneuver. This subsystem calculates constraints onthe acceleration based on the current power and available power of theaircraft.

The engine power (or torque) is measured and filtered. The maximum poweravailable from the engines is typically a function of ambientconditions. Helicopter turbine engines have less power available at highaltitude or high ambient temperatures. In some cases the power may belimited by transmission limits. A schedule is used to determine maximumpower. Maximum power is compared to the measured power to determine thepower margin. For the deceleration limit, the power is compared to theminimum power, which would be set to 0 or some small value to preventover speed of the rotor and drive system.

A simple relationship is used to estimate acceleration limits from thepower margin. The kinetic energy of the aircraft is represented by:

$\begin{matrix}{{K.E.} = {\frac{1}{2}\mspace{14mu}{mV}^{2}}} & {{Eqn}.\mspace{14mu} 13}\end{matrix}$

The power increment required to accelerate can be estimated by takingthe derivative of the kinetic energy equation and also including anefficiency factor, □.

$\begin{matrix}{{\Delta\; P} = \frac{{mV}\overset{.}{V}}{\eta}} & {{Eqn}.\mspace{14mu} 14}\end{matrix}$

Substituting the power margin for □P and allowing for conversion fromhorsepower:

$\begin{matrix}{{\overset{.}{V}}_{\lim} = {550\frac{{\eta\Delta}\;{HP}}{mV}}} & {{Eqn}.\mspace{14mu} 15}\end{matrix}$

When implementing equation 15, it is necessary to put a lower limit onthe speed term to avoid division by zero.

The relationship expressed in equation 15 provides an approximateexpression for the acceleration limits. If the aircraft is acceleratingat a limit defined by Equation 15, the power margin may not approachzero in steady-state. This may allow the aircraft to exceed the powerlimit in steady-state or operate too conservatively below the powerlimit. If the aircraft is operating at an acceleration limit, acorrection mechanism is engaged. The power margin is passed though anintegrator compensator and added to the acceleration limit. This stepforces the power margin to approach zero if the aircraft is operating atan acceleration limit.

H. Lateral/Directional Control

FIG. 10 is a schematic diagram of the coupled lateral/directionalcontrol subsystem 74. The operation of the lateral/directional controlsubsystem 74 is directly analogous to the operation of thelongitudinal/vertical control subsystem 72 illustrates by FIG. 7 anddiscussed above. The purpose of the lateral/directional controlsubsystem 74 is to integrate the redundant controls of the compoundaircraft 2 for roll and yaw. Those redundant controls include the 7cyclic, differential flaperons, sector, rudder and propeller pitch.

The inputs to this subsystem are the commanded roll attitude, thecommanded yaw rate, and the trim roll attitude. The outputs of thesystem are the lateral cyclic, differential flaperons, and yaw control.The lateral/directional control system 74 uses a model following/modelinversion architecture. The model and the control subsystem 74 areconfigured to achieve attitude command/attitude hold in roll so that theaircraft will respond to a command to achieve a desired roll and rollrate and will hold a selected roll angle. The model and the controlsubsystem 74 also are configured to achieve rate command heading/headinghold in yaw, so that the aircraft will achieve a commanded yaw rate andwill hold a specified yaw angle in flight. The model and control systemcan be configured to exhibit any desired dynamic response, such as therequirements of the ADS-33 specification set. The desired roll responseis second order and can be designed to meet the ADS-33 roll bandwidthrequirements. The desired yaw response is first order can be designed tomeet the ADS-33 yaw bandwidth requirements.

The pilot commands 6 to the lateral/directional subsystem 74 include thepilot-commanded roll rate (ΔΦ_(CMD)) and the pilot-commanded yaw rate(τ_(CMD)). The subsystem 74 also receives the trim roll attitude fromthe optimum trim schedule 68. The aircraft condition information 66inputs include the roll attitude and roll rate, yaw attitude, airspeed,and pitch attitude. The outputs of the system are supplied to the rotormixing module 84, the wing mixing module 82 and the VTDP mixing module86, as illustrated by FIG. 6. As in Equations 8 to 10, a weightedpseudo-inverse is used to distribute control among the threelateral-directional controls.

I. Turn Coordination/Yaw Rate Command

FIG. 11 illustrates in detail the turn coordination/yaw rate commandsubsystem 76 shown on FIG. 6. The turn coordination/yaw rate commandsubsystem 76 is desirable because the turning characteristics andrequirements of the compound aircraft 2 flying at low speed aredifferent from the turning characteristics and requirements when theaircraft 2 is flying at high speed. The Yaw Rate Command mode isdesirable for speeds below 50 knots and the Turn Coordination mode isdesirable for speeds above 60 knots. A simple blending scheme is usedfor transitions between 50 and 60 knots. At low speed (below 50 knots)the turn coordination/yaw rate control system 76 will allow the aircraft2 to turn at a yaw rate (r) of up to 60 degrees/second. At high speed(greater than 60 knots) the control system allows a maximum lateralacceleration of 20 ft/sec². For speeds between 50 and 60 knots, the yawrate and lateral acceleration commands are blended. The output of theturn coordination/yaw rate command subsystem 76 (r_(CMD)) is feddirectly to the lateral/directional control subsystem 74, as shown byFIG. 6.

The turn coordination mode calculates an effective yaw rate for turncoordination. This yaw rate can be fed directly to theLateral/Directional AFCS, so no changes to this AFCS need to be made asthe aircraft transitions from low speed to high speed mode. Thecommanded lateral acceleration is compared to measured value. The turncoordination controller can be derived from the equation of motion forthe lateral body velocity:{dot over (v)}=a _(y) −ur+pw+g sin φ cos θ  Eqn. 16With some simplifying assumptions one can derive the control law:

$\begin{matrix}{r_{cmd} = \frac{{G_{TC}\left( {a_{y_{cmd}} - a_{y}} \right)} + {g\;\sin\;{\phi cos\theta}}}{V}} & {{Eqn}.\mspace{14mu} 17}\end{matrix}$The yaw rate command is calculated to correct the lateral accelerationerror and then fed to the Lateral-Directional AFCS.J. Return to Trimmed Flight after a Maneuver

At the end of a maneuver, the compound aircraft 2 returns automaticallyto trimmed flight optimized to achieve the selected overall operationalobjective. The microprocessor 38 constantly updates the combination oftrim settings selected from the trim schedule based on feedback in theform of the measured condition of the compound aircraft 2 as detected bysensors 40. The selected combination of trim settings changes relativelyslowly because the parameters on which the selection is based(parameters such as airspeed, altitude and vehicle gross weight) changerelatively slowly.

During a maneuver of the compound aircraft 2, the dynamic inversioncontroller is constantly receiving feedback and is adding or subtractingcontrol effector deflections to achieve the commanded maneuver. Thecorrections of the dynamic inversion controller are based mainly on thefeedback of aircraft condition variables that change rapidly with time,for example angular rate and aircraft attitude.

When the compound aircraft 2 reaches equilibrium after a maneuver, theaircraft conditions as detected by the sensors 40 (and especially thefast parameters such as aircraft attitude) will have reached nearsteady-state and the feedback signals will approach zero or a smallvalue. In addition, once the aircraft 2 is in trimmed flight, the pilotwill have moved the control inceptors back to the detent position,indicating that the pilot is commanding trimmed flight. The controlsystem 62 will adjust the control effectors consistent with theconstantly-updated combination of trim settings to achieve the selectedoverall operational objective.

K. Simulation Study

The benefits of a compound aircraft 2, including a compound aircraft 2having the control system 62 of the Invention, have been demonstrated. Asimulation study was performed that compared the performance of an AH-60Blackhawk helicopter to a similar Blackhawk helicopter equipped as acompound aircraft 2. The simulation study evaluated conditions ofminimum power usage and minimum vibration, among others, in trimmedflight for comparable aircraft with comparable loads and for comparablealtitudes and air temperatures.

The simulation study showed that the compound aircraft 2 was capable ofhigher speeds than the comparable helicopter and that the compoundaircraft 2 was capable of operating with less power than the helicopterfor any particular airspeed. When optimized for minimum vibration, thesimulated compound aircraft 2 offered significant vibration reductioncompared to the helicopter. The simulation study showed that reducedvibration did not necessarily coincide with reduced power.

The simulation study demonstrates that no one set of control settings isoptimal for all possible missions of the compound aircraft 2 and thatcontrol settings 78 may be optimized to accomplish an overall objective,such as minimizing vibration or minimizing power. The study alsodemonstrated that the optimum control settings 78 to accomplish anobjective vary with the flight condition of the aircraft 2, such asairspeed.

L. Control of Maneuvering Flight

The control system 62 of the Invention may be used to achieve auser-selected overall operational objective for maneuvering flight aswell as trimmed flight. For maneuvering flight, the objective ofachieving a selectable overall operational objective may be accomplishedby selecting appropriate weighting factors for distribution of controlamong the various control effectors.

An investigation conducted concerning the Invention demonstrated thatselection of weighting factors can limit structural loads duringcritical maneuvers. For example, the weighting factor “K_(θ)” is appliedin the Speed Command/Speed Holding subsystem to allocate forward thrustbetween the main rotor and the propeller. Different values of the K_(θ)parameter may be scheduled and stored in computer memory 60. Theappropriate value for K_(θ) may be selected by the microprocessor 38 toachieve a selectable overall operational objective based on the currentcondition of the aircraft and on the control inceptor input.

Other weighting factors applicable to the other control effectors appearin the model inversions of FIGS. 6-11 and are indicated collectively inthe figures and the equations by the symbol “w.” In the discussion abovefor longitudinal control and vertical control, the weighting factors areindicated as “w_(long), w_(coll), w_(e), and w_(δ) _(F) ” for weightingfactors for longitudinal cyclic, collective, elevator and flaperons,respectively. Similar weighting factors exist for theLateral/Directional system; namely, w_(lat), w_(dir), w_(df), relatingto lateral cyclic, directional control and differential flaperons,respectively. Other weighting factors also may be applied.

The weighting factors may be selected by the microprocessor 38 from aweighting factor schedule stored in computer memory 60, just as acombination of control settings for trim is selected from the trimschedule. The weighting factor schedule may be configured to selectparticular weights based on an overall objective to be achieved (i.e.,vibration reduction, life cycle cost reduction, prevention of overstressing a component) and upon current flight conditions such asairspeed, altitude, vehicle gross weight and center of gravity position.Other parameters may be used to select the appropriate weighing factorfrom the schedule.

M. Other Applications

The control system 62 of the Invention may be applied in any situationwhere an objective can be accomplished through any one of a plurality ofcombinations of control variables and where the different combinationsof control variables achieve different overall objectives. In thecontrol system 62 of the Invention, the user selects among the overallobjectives. The control system 62 then selects the appropriatecombination of control variables to accomplish the objective consonantwith the selected overall objective.

A user may actuate a selector 41 to select an overall operationalobjective by any conventional means, such as by selecting an icon on acomputer display, by throwing a switch, or by affixing a jumper to acircuit board.

In describing the above embodiments of the invention, specificterminology was selected for the sake of clarity. However, the inventionis not intended to be limited to the specific terms so selected, and itis to be understood that each specific term includes all technicalequivalents that operate in a similar manner to accomplish a similarpurpose.

We claim:
 1. A compound aircraft, the compound aircraft comprising: a. afuselage; b. a rotor, said rotor being rotatably attached to saidfuselage; c. a wing, said wing being attached to said fuselage; d. aplurality of control effectors attached to said fuselage; e. amicroprocessor operably connected to said plurality of controleffectors, said microprocessor being configured to operate said controleffectors; f. a computer memory operably connected to saidmicroprocessor; g. a plurality of selectable overall operationalobjectives stored within said computer memory; h. a selector, saidselector being configured to select a one of said overall operationalobjectives from among said plurality of overall operational objectives,said selector being operably connected to said microprocessor; i. asensor, said sensor being configured to detect a condition of thecompound aircraft, said sensor being operably connected to saidmicroprocessor; j. a plurality of combinations of trim control settingsto control at least one of a roll, a pitch and a yaw of the compoundaircraft, said plurality of combinations of trim control settings beingselectable by said microprocessor; k. said microprocessor beingconfigured to receive a command, each of said plurality of trim controlsettings being configured to achieve said command when applied by saidmicroprocessor, a one of said plurality of trim control settingscorresponding to said selected operational objective, saidmicroprocessor being configured to select said one of said plurality oftrim control settings corresponding to said selected operationalobjective, said microprocessor being configured to operate said controleffectors to implement said selected one of said plurality ofcombinations of trim control settings.
 2. The compound aircraft of claim1 wherein at least two of said control effectors are configured tocontrol said roll of the compound aircraft and wherein said selected oneof said plurality of said combinations of trim control settings includessaid trim control settings for said at least two of said controleffectors configured to control said roll of the compound aircraft. 3.The compound aircraft of claim 2, the compound aircraft furthercomprising: a plurality of combinations of weighting factors, saidplurality of combinations of said weighting factors being stored in saidcomputer memory, said microprocessor being programmed to select a one ofsaid plurality of said combinations of weighting factors in response tosaid command, said selected one of said plurality of said combinationsof weighting factors being associated with said selected one of saidoverall operational objectives, said selected one of said plurality ofsaid combinations of weighting factors including said weighting factorsto control said roll of the compound aircraft in maneuvering flightutilizing at least said two of said control effectors configured tocontrol said roll of the compound aircraft.
 4. The compound aircraft ofclaim 1 wherein at least two of said control effectors are configured tocontrol said pitch of the compound aircraft, said selected one of saidplurality of said combinations of trim control settings including saidtrim control settings for said at least said two of said controleffectors to control said pitch of the compound aircraft.
 5. Thecompound aircraft of claim 4, the compound aircraft further comprising:a plurality of combinations of weighting factors, said plurality ofcombinations of said weighting factors being stored in said computermemory, said microprocessor being programmed to select a one of saidplurality of said combinations of weighting factors in response to saidcommand, said selected one of said plurality of said combinations ofweighting factors being associated with said selected one of saidoverall operational objectives, said selected one of said plurality ofsaid combinations of weighting factors including said weighting factorsto control said pitch of the compound aircraft in maneuvering flightutilizing at least said two of said control effectors configured tocontrol said pitch of the compound aircraft.
 6. The compound aircraft ofclaim 1 wherein at least two of said control effectors are configured tocontrol said yaw of the compound aircraft, said selected one of saidplurality of combinations of trim control settings including trimcontrol settings for said at least two of said control effectorsconfigured to control said yaw of the compound aircraft.
 7. The compoundaircraft of claim 6, the compound aircraft further comprising: aplurality of combinations of weighting factors, said plurality ofcombinations of said weighting factors being stored in said computermemory, said microprocessor being programmed to select a one of saidplurality of said combinations of weighting factors in response to saidcommand, said selected one of said plurality of said combinations ofweighting factors being associated with said selected one of saidoverall operational objectives, said selected one of said plurality ofsaid combinations of weighting factors including said weighting factorsto control said yaw of the compound aircraft in maneuvering flightutilizing said at least two of said control effectors configured tocontrol said yaw.
 8. A compound aircraft, the compound aircraftcomprising: a. a fuselage; b. a rotor, said rotor being rotatablyattached to said fuselage; c. a wing, said wing being attached to saidfuselage; d. a plurality of control effectors attached to said fuselage;e. a microprocessor operably connected to said plurality of controleffectors; f. a sensor, said sensor being configured to detect acondition of the compound aircraft, said sensor being operably connectedto said microprocessor; g. a computer memory operably connected to saidmicroprocessor; h. said microprocessor being configured to receive acommand; i. said microprocessor being configured to operate saidplurality of control effectors to apply a combination of trim controlsettings that will achieve said command in light of said condition ofthe compound aircraft; j. a selector, said selector being configured toselect an overall operational objectives from among a plurality ofoverall operational objectives, said selector being operably connectedto said microprocessor; k. said combination of trim control settingsbeing a one of a plurality of combinations of trim control settings eachof which would achieve said command in light of said condition of thecompound aircraft, said microprocessor being configured to apply saidone combination of trim control settings that is consistent with saidselected operational objective.
 9. The compound aircraft of claim 8, thecompound aircraft further comprising: a control inceptor, said controlinceptor being operably attached to said microprocessor, said controlinceptor being configured to receive said command and to communicatesaid command to said microprocessor.
 10. The compound aircraft of claim9 wherein said control inceptor is configured to receive said commandfrom a human being.
 11. The compound aircraft of claim 10 wherein saidplurality of combinations of trim control settings is stored in saidcomputer memory and wherein said plurality of overall operationalobjectives is stored in said computer memory, said microprocessor beingconfigured to select said one of said plurality of combinations of trimcontrol settings from among said plurality of combinations of trimcontrol settings stored in said computer memory.
 12. The compoundaircraft of claim 11 wherein said plurality of control effectors areconfigured to control at least a one of a roll, a pitch, and a yaw ofsaid fuselage.
 13. The compound aircraft of claim 8, the compoundaircraft further comprising: a plurality of combinations of weightingfactors configured to be applied by said microprocessor to control thecompound aircraft in maneuvering flight, said plurality of combinationsof weighting factors being stored in said computer memory, saidmicroprocessor being configured to select a one of said plurality ofcombinations of weighting factors in response to said command, saidselected one of said plurality of combinations of weighting factorsbeing consistent with said selected one of said overall operationalobjectives.
 14. The compound aircraft of claim 13 wherein said selectedone of said plurality of said combinations of weighting factors includessaid weighting factors to control at least a one of said roll, saidpitch and said yaw of the compound aircraft.
 15. The compound aircraftof claim 14 wherein said plurality of combinations of trim controlsettings is stored in said computer memory and wherein said plurality ofoverall operational objectives is stored in said computer memory, saidmicroprocessor being configured to select said one of said plurality ofcombinations of trim control settings from among said plurality ofcombinations of trim control settings stored in said computer memory.